Method for operating a lean premix burner of an aircraft gas turbine and device for carrying out the method

ABSTRACT

The present invention relates to a method for operating a lean premix burner of an aircraft gas turbine, where fuel and primary supporting air are supplied by means of a supporting burner (pilot burner) arranged centrically to the burner axis, where secondary air surrounding the supporting burner is supplied, and where fuel and air are supplied by means of a main burner, characterized in that the primary supporting air is supplied in an amount of 5 vol % to 10 vol % of the total air quantity, that the secondary supporting air is supplied in an amount of 5 vol % to 20 vol % and that 35 vol % to 75 vol % of the total air quantity are supplied via the main burner in the partial load range and in the full load range.

This invention relates to a method for operating a lean premix burner ofan aircraft gas turbine, where fuel and primary supporting air aresupplied by means of a supporting burner (pilot burner) arrangedcentrically to the burner axis and where fuel and air are supplied bymeans of a main burner.

It is known from the state of the art to use two fuel atomizers, i.e. asupporting burner and a main burner, in lean premix burners. Thesupporting burner is arranged centrically in the main burner. Thesupporting burner is here usually designed as a pressure swirl atomizer.The lean premix burner includes here two fuel lines for supplying thesupporting burner and the main burner. In operation, the supportingburner is used for igniting the aircraft gas-turbine engine and in lowload conditions, whereas the main burner is put into operation atpartial load and is used up to maximum load. The supporting burner isdesigned here for the ignition operation and for a stable combustionduring the engine starting phase.

The state of the art is described in the following in light of FIG. 2,which shows a burner 32 arranged on a combustion chamber head 31 andsupplying the combustion chamber with fuel and approximately 10 vol % to20 vol % of the total air (strictly speaking this is mass %, but in thiscase they are identical since the air has a constant temperature.) As aresult, a rich zone 33 is formed, which is arranged directly downstreamof the burner 32. A further 30 vol % to 40 vol % of air are suppliedthrough mixing air openings 34 to 37. This results in an air admixtureto the rich flame in the flame zone 38. Downstream of this flame zone38, a lean zone 39 is provided. The remaining air of 40 vol % to 50 vol% is used for cooling and flows through an inner combustion chamber wall40 and an outer combustion chamber wall 41, which maintain the flame.FIG. 2 thus shows a standard burner with a rich zone supplied with airand followed by a lean zone.

FIG. 3 shows an embodiment according to the state of the art, where theburner 32 includes means for mixing air and fuel. A direct or furtherflame zone 38 as shown in FIG. 2 can be dispensed with. The entireburner (supporting burner and main burner) passes 50 vol % to 80 vol %of the total air into the combustion chamber. The remaining air quantityof 20 vol % to 50 vol % is used for cooling. The burner includes twofuel circuits and thus permits the supply of fuel through two concentricfuel atomizers. The supporting burner 42 with the associated atomizersupplies 5 vol % to 15 vol % of the total air and creates a small richzone 33 which is used for starting the engine and for flame stability.The concentric main fuel atomizer 43 supplies fuel at medium to maximumload conditions and 40 vol % to 75 vol % of the total air 44. Thiscreates a lean zone 39 surrounding the rich zone 33. This lean zone 39is responsible for low pollutant emissions, in particular of NOx.

The main drawback of the solution shown in FIG. 2 is that high pollutantemissions result, in particular of NOx, and that in high load conditionssoot is emitted, since the combustion conditions approximate to astoichiometric or rich combustion state. It was therefore attempted inthe solution described in FIG. 3 to optimize the combustion process by alean burner concept. This however has the disadvantage that thesupporting burner (pilot burner) has a reduced flame stability at lowoutput of the aircraft gas turbine. In medium load conditions, thecombustion by the main burner is too lean to operate effectively,leading to increased fuel consumption by an aircraft. Additionally, theabsence of high air admixing results in low oxidation of soot, so thatconsiderable soot quantities are emitted from the aircraft gas turbineat medium load.

The object underlying the present invention is to provide a method—and adevice for carrying out the method—for operating a lean premix burnerwhich avoid the disadvantages of the state of the art and enable, inparticular, a good, stable and low-pollutant combustion.

It is a particular object of the present invention to provide solutionto the above problematics by a combination of the features of theindependent Claims. Further advantageous embodiments of the presentinvention become apparent from the sub-claims.

It is thus provided in accordance with the invention that between therich zone and the lean zone an intermediate admixing zone with a highjet-like admixture of air is formed, into which zone an additionalfuel/air flow is introduced. This results in optimized combustion inpartial load areas too, and has the advantage that the soot emissionsare reduced by improved oxidation of the soot. Furthermore, there isimproved combustion with better efficiency, since the very lean zonesknown from the state of the art are avoided. Due to the intermediateadmixing zone, a combustion zone is created which is closer tostoichiometric fuel/air ratios. Although this zone is still lean, itavoids the disadvantages of a too-lean combustion zone.

In accordance with the invention, there is an enrichment of the richzone with a lower air proportion by the supporting burner. Instead, theadditional air is introduced into the intermediate admixing zone. Thisleads to good flame stability and good ignitability of the aircraft gasturbine.

The present invention is described in the following in light of theaccompanying drawing, showing exemplary embodiments. In the drawing,

FIG. 1 shows a schematic representation of a gas-turbine engine inaccordance with the present invention,

FIG. 2 shows a longitudinal sectional view of a combustion chamber inaccordance with the state of the art,

FIG. 3 shows a schematic representation of a further variant of acombustion chamber in accordance with the state of the art by analogywith FIG. 2,

FIG. 4 shows a simplified longitudinal sectional view of a combustionchamber in accordance with a first exemplary embodiment of the inventionby analogy with the representation of FIG. 3,

FIG. 5 shows a representation of a further exemplary embodiment byanalogy with FIG. 4,

FIG. 6 shows a sectional view of a further exemplary embodiment byanalogy with FIGS. 4 and 5,

FIG. 7 shows an enlarged partial representation of the flow conditionsof the inventive solution in accordance with FIG. 6,

FIGS. 8 to 11 show sectional, front and perspective views of differingexemplary embodiments of flame stabilizers and secondary air recesses,

FIG. 12 shows a graphic representation of the equivalence ratio as afunction of the thrust in accordance with the state of the art,

FIG. 13 shows a graphic representation of a lean premix burner, byanalogy with FIG. 12, and

FIG. 14 shows a representation of the inventive solution by analogy withFIGS. 12 and 13.

The gas-turbine engine 10 in accordance with FIG. 1 is a generallyrepresented example of a turbomachine where the invention can be used.The engine 10 is of conventional design and includes in the flowdirection, one behind the other, an air inlet 11, a fan 12 rotatinginside a casing, an intermediate-pressure compressor 13, a high-pressurecompressor 14, a combustion chamber 15, a high-pressure turbine 16, anintermediate-pressure turbine 17 and a low-pressure turbine 18 as wellas an exhaust nozzle 19, all of which being arranged about a centerengine axis 1.

The intermediate-pressure compressor 13 and the high-pressure compressor14 each include several stages, of which each has an arrangementextending in the circumferential direction of fixed and stationary guidevanes 20, generally referred to as stator vanes and projecting radiallyinwards from the engine casing 21 in an annular flow duct through thecompressors 13, 14. The compressors furthermore have an arrangement ofcompressor rotor blades 22 which project radially outwards from arotatable drum or disk 26 linked to hubs 27 of the high-pressure turbine16 or the intermediate-pressure turbine 17, respectively.

The turbine sections 16, 17, 18 have similar stages, including anarrangement of fixed stator vanes 23 projecting radially inwards fromthe casing 21 into the annular flow duct through the turbines 16, 17,18, and a subsequent arrangement of turbine blades 24 projectingoutwards from a rotatable hub 27. The compressor drum or compressor disk26 and the blades 22 arranged thereon, as well as the turbine rotor hub27 and the turbine rotor blades 24 arranged thereon rotate about theengine axis 1 during operation.

FIG. 4 shows in a schematic representation a longitudinal sectional viewof a burner in accordance with the invention. It includes a supportingburner 42 and a fuel atomizer 43 surrounding the latter, both formingpart of a burner 32 mounted on a combustion chamber head 31. Thecombustion chamber includes an inner combustion chamber wall 40 and anouter combustion chamber wall 41. Air and fuel are supplied by thesupporting burner 42 for forming a rich zone 33 immediately adjoiningthe supporting burner 42. In total, an air quantity of approx. 50 vol %to 80 vol % of the total air is supplied to the combustion chamber bythe burner. The fuel is supplied via two concentric atomizers. Only asmall amount of air (5 vol % to 10 vol % of the total combustion chamberair) is supplied via the atomizer of the supporting burner 42, thusensuring that the rich zone 33 is created.

The rich zone 33 is delimited and partially enclosed by an intermediateadmixing zone 45. An air quantity of 5 vol % to 20 vol % of the totalcombustion air of the combustion chamber is introduced into theintermediate zone in order to provide a second zone or secondarysupporting zone (intermediate admixing zone) 45 forming a furtheradmixing zone (quenching zone).

The main fuel atomizer 43 supplies fuel and air. The air quantitysupplied is 35 vol % to 75 vol % of the total combustion chamber air.The main fuel atomizer 43 is used during medium to maximum load statesof the aircraft gas turbine. By supplying air and fuel via the main fuelatomizer 43, a lean zone 39 is created which surrounds the intermediateadmixing zone 45 and adjoins the latter in the axial direction (flowdirection).

FIG. 6 shows a detailed representation of a further exemplary embodimentof the invention, by analogy with the representation in FIG. 4.Identical parts are provided with the same reference numerals, as is thecase in the following exemplary embodiments.

FIG. 5 shows in detail the supporting burner 42 with a fuel outlet 47.The supporting burner is concentrically surrounded by an annular airpassage 48, in which a swirler element 49 is arranged. The escapingair/fuel mixture creates the rich zone 33.

The intermediate admixing zone 45 is formed by the further supply of airand fuel. A concentric annulus 50 is provided for this. The designpermits a greater pressure drop, in order to generate higher airvelocities at the place where air is introduced into the combustionchamber. This results in good mixing with the rich zone 33. Thesecondary air supply 51, 52 and 53 can take place through suitablerecesses described in the following in conjunction with FIGS. 8 to 10.

The main fuel is supplied through a concentric main air supply 54 andatomized by the inner air supply 55 and mixed with the latter. A swirlis imparted by an inner main swirler element 56. The main fuel is alsoguided through an outer air supply 57 and atomized and mixed with it,with a swirl being imparted to this air supply by means of an outer mainswirler element 58. The flame resulting from the main burner surroundsthe intermediate admixing zone 45 and forms a lean zone 39.

The secondary air can be supplied at different points (secondary airsupply 51, 52 or 53). This supply can take place singly or incombination.

FIG. 11 shows an axial longitudinal sectional view plus a front view ofan exemplary embodiment of a burner in accordance with the invention. Itis shown here that the secondary air recesses 52 can be designed in theform of round holes provided on a flame stabilizer.

In the exemplary embodiment in FIG. 9, the secondary air is supplied bytubes (chutes) provided on the flame stabilizer 59. It can be suppliedin either an axial or a tangential alignment in order to impart a swirlto the secondary air. Between 4 and 36 of these outlet tubes (chutes)can be provided, being at angles of 0° and 60° to the burner axis.

FIG. 11 shows a further design variant in which the secondary airrecesses 52 are designed in the form of slots. Between 4 and 36 slotscan be provided, and can have an angle between 0° and 60° relative tothe burner axis in order to impart an additional swirl to the air.

FIG. 8 shows a further exemplary embodiment with V-shaped slots 52,which can also be provided in a number between 4 and 36. Here too it ispossible to incline the V-shaped slots relative to the burner axisbetween 0° and 60° for further swirling of the air.

The burner described above can also be designed with an onflowsupporting burner, as is shown in FIG. 6. The supporting fuel issupplied through a fuel outlet 47. The supporting air is suppliedthrough an inner air passage 48 with a swirler element 49 and an outerannulus 50 with a swirler element 56.

FIG. 7 shows that in accordance with the invention a second supportingflame stabilization zone Y is formed additionally to zone X and to therich zone 33. The zone Y leads to an improved interaction between thesupporting burner and the main burner. In certain operating conditions,the main flame can also be stabilized in zone Y.

In accordance with the invention, an additional intermediate zone isthus created by which combustion in the combustion chamber can takeplace in a controlled and optimized way. This leads to the supportingburner zone being able to operate in a stable manner, without any fearof the supporting burner being extinguished. The intermediate zone canbe operated even in relatively high load conditions without sootemissions. Furthermore, the intermediate admixing zone improvescombustion efficiency (total combustion) during staged operation of themain burner. This leads to a minimum drop in the efficiency ofcombustion during a transition from operation of the supporting burnerto combined operation of the supporting burner and of the main burner.

In the following, the invention is again explained in respect of themethod in accordance with the invention in light of FIG. 14, where theillustrations in FIGS. 12 and 13 reflect the underlying state of theart.

FIG. 12 shows a diagram in which the thrust is plotted in percentages asa function of the equivalence ratio between air and fuel. With anequivalence ratio of 1, there is a stoichiometric ratio, below 1 to 0the result is a rich combustion, while above 1 a lean combustion isobtained. These illustrations are also shown in FIGS. 13 and 14.

Furthermore, the information for the combustion zones relates to FIGS. 2to 5.

FIG. 12 shows an illustration from the state of the art which has highemission values. In particular, at high thrust or high output,respectively, the NOx values are high and there is a lot of soot. Therespective equivalence ratios of the individual zones achieve here, asis indicated in FIG. 12 for the zones 33, 38 and 39, values having anequivalence ratio of 1 or a rich equivalence ratio close to thestoichiometric value. By contrast, the result for the supporting burneris good flame stability.

To avoid the drawbacks of increased soot formation and high NOxemissions, solutions were proposed as shown in FIG. 13. While FIG. 12 inparticular relates to the representation in FIG. 2, the values shown inFIG. 13 are based especially on an embodiment in accordance with FIG. 3.As shown in FIG. 13, the supporting burner is operated somewhat moreleanly. This leads to a good combustion, but at the same time generatesa lot of soot. At the same time a reduced stability at low load resultsfrom the leaner supporting burner. As also shown in FIG. 13, the burneris set very lean at medium thrust, so that in this partial load area ortransition area there is no good combustion in particular in zone 39. Alow efficiency thus applies, and this leads to increased fuelconsumption of an aircraft.

Furthermore, the absence of the flame zone 38 and the poor interactionbetween the mixing air 36 lead to a poor oxidation of soot, resulting inhigh soot emissions.

The drawbacks of the mode of operation shown in FIG. 13 can be reducedaccording to the state of the art in that the supporting burner isenlarged to pass a larger air quantity through the supporting burner.Soot formation could be reduced by this, but this has the negativeeffect that higher NOx emissions result. Moreover, a leaner operation ofthe supporting burner leads to a lower stability. Furthermore, a secondsupporting burner circuit with a total of three fuel circuits could beintroduced, but this would increase the complexity of the overallsystem, involving additional costs for fuel injection nozzles, fuelsystems and control systems.

Based on the procedures described above, a completely different solutionwas created in accordance with the invention, and is explained in lightof FIG. 14.

The solution in accordance with the invention was described above inparticular in conjunction with the design solution according to FIG. 4.

In accordance with the invention, a secondary supporting zone orintermediate admixing zone 45 is formed, as explained above, which isachieved by diverting air/fuel from the rich zone 33. Furthermore, thereis a diversion of fuel and air from the total airflow 44. By doing so,an additional flow 46 is used, as is shown in FIG. 4.

The solution in accordance with the invention results in the followingadvantages:

As shown in FIGS. 4 and 5, the addition of the secondary supportingzone/intermediate admixing zone 45 leads to a reduction in the sootemissions, caused by an improved oxidation of the soot. Furthermore,there is an improved combustion efficiency due to the reduction of verylean zones. The zone of the main burner remains lean, but zone 45 formsa secondary supporting zone or intermediate admixing zone which iscloser to the stoichiometric fuel/air ratio. Furthermore, the zone 45leads to a reduction in soot formation. The zone 33 (zone of thesupporting burner 42) can be operated with a richer fuel/air mixturethan in the solutions known from the state of the art. This leads to animproved flame stability.

As explained above, the core of the invention is the additionalintroduction of a secondary supporting zone or intermediate admixingzone 45. This leads to the supporting burner being capable of operationwith a constant combustion zone, thereby ensuring stable operation andpreventing the flame from being extinguished (flame-out). Both thesupporting burner zone and the secondary supporting zone/intermediateadmixing zone 45 can be operated without problems resulting with regardto soot emissions. Furthermore, the secondary supportingzone/intermediate admixing zone 45 improves the efficiency of combustionduring a staged operation of the main burner. This leads to a minimumreduction in the combustion efficiency during the transition fromoperation with the supporting burner to combined operation of thesupporting burner and of the main burner.

LIST OF REFERENCE NUMERALS

-   1 Engine axis-   10 Gas-turbine engine/core engine-   11 Air inlet-   12 Fan-   13 Intermediate-pressure compressor (compressor)-   14 High-pressure compressor-   15 Combustion chamber-   16 High-pressure turbine-   17 Intermediate-pressure turbine-   18 Low-pressure turbine-   19 Exhaust nozzle-   20 Guide vanes-   21 Engine casing-   22 Compressor rotor blades-   23 Stator vanes-   24 Turbine blades-   26 Compressor drum or disk-   27 Turbine rotor hub-   28 Exhaust cone-   31 Combustion chamber head-   32 Burner-   33 Rich zone-   34, 35, 36, 37 Mixing air-   38 Flame zone-   39 Lean zone-   40 Inner combustion chamber wall-   41 Outer combustion chamber wall-   42 Supporting burner-   43 Fuel atomizer-   44 Total air-   45 Secondary supporting zone/intermediate admixing zone-   46 Additional flow-   47 Fuel outlet-   48 Air passage-   49 Swirler element-   50 Annulus-   51, 52, 53 Secondary air supply/secondary air recesses-   54 Concentric main air supply-   55 Inner air supply of main burner-   56 Inner main swirler element-   57 Outer air supply of main burner-   58 Outer main swirler element-   59 Flame stabilizer-   60 Supporting air supply

1. Method for operating a lean premix burner of an aircraft gas turbine,where fuel and primary supporting air are supplied by means of asupporting burner (pilot burner) arranged centrically to the burneraxis, where secondary air surrounding the supporting burner is supplied,and where fuel and air are supplied by means of a main burner,characterized in that the primary supporting air is supplied in anamount of 5 vol % to 10 vol % of the total air quantity, that thesecondary supporting air is supplied in an amount of 5 vol % to 20 vol %and that 35 vol % to 75 vol % of the total air quantity are supplied viathe main burner in the partial load range and in the full load range. 2.Method in accordance with claim 1, characterized in that adjacent to thesupporting burner a rich zone is formed, that the rich zone is enclosedby an intermediate admixing zone, that the intermediate admixing zone isenclosed and that the intermediate admixing zone is enclosed by a leanzone.
 3. Aircraft gas turbine lean premix burner for carrying out themethod in accordance with claim 1, characterized in that a flamestabilizer concentrically surrounding the supporting burner is providedwith secondary air recesses.
 4. Premix burner in accordance with claim3, characterized in that the secondary air supply recesses are providedin the form of straight or V-shaped slots.
 5. Premix burner inaccordance with claim 3, characterized in that the secondary air supplyrecesses are provided in the form of tubes (chutes).